1. Field of Invention
This invention relates to systems for controlling the attitude of a spacecraft. Specifically, the present invention relates systems for controlling the attitude of spinning satellites.
2. Description of the Related Art
Satellites are used in a variety of demanding applications ranging from communications systems and global positioning systems to space telescope systems. Such applications often require satellites with accurate maneuvering, station keeping, and attitude adjustment capabilities.
Systems for accurately adjusting satellite attitude are particularly important in communications and space telescope applications. For example, ground based receivers or transceivers aimed at a communications satellite often require the satellite to remain at its current position and orientation to communicate with the satellite. Similarly, space telescope satellites are often precisely oriented for focusing on a particular region in space.
Without attitude control mechanisms and spacecraft maneuvering systems, a satellite will spin about a random axis. The spin may result from moments generated during satellite launch, from solar radiation pressure, and/or gravity gradients.
To control spacecraft attitude, thrusters are often employed. The thrusters provide coarse control over the orientation of the satellite but typically lack the ability to make fine attitude adjustments required by many satellite applications. The thrusters also consume excess fuel, which increases the weight and cost of the spacecraft and limits the usable life span of the spacecraft.
The thrusters are typically employed on most spacecraft, including those with non-zero initial momentum such as spinning communications satellites. When used on a spinning spacecraft, the thrusters must provide large torques to move the spacecraft momentum vector to a desired location. (This is known as the bicycle wheel effect.) The requisite torques result in additional fuel consumption.
To provide finer control over spacecraft orientation and to reduce the need for expensive thrusters, a system employing a gimbaled momentum wheel is often employed. The system is disclosed in U.S. Pat. No. 5,441,222, by H. Rosen, entitled ATTITUDE CONTROL OF A SPINNING SPACECRAFT, the teachings of which are herein incorporated by reference. The system includes a two-axis gimbaled momentum wheel. The wheel is spun to cancel existing spacecraft momentum, resulting in a zero momentum spacecraft. The zero momentum spacecraft is then oriented via internal moments created by actuators that push on the wheel assembly to create reaction forces in desired directions.
Use of the gimbaled momentum wheel however, has several drawbacks. Spacecraft design constraints typically limit the size and mass of the momentum wheel. Consequently, the momentum wheel must spin at high rates to cancel the total momentum of the spacecraft. Often the momentum wheel must spin faster than 5000 revolutions per minute. This can result in undesirable high frequency spacecraft vibrations that are difficult to remove. Expensive and complex momentum wheel isolation systems and/or control loops are often required to reduce the undesirable high frequency disturbances.
Hence, a need exists in the art for a cost-effective system for precisely orienting a spacecraft. There is a further need for a system that does not introduce high frequency spacecraft vibrations or consume excess fuel.
The need in the art is addressed by the attitude control system for a spacecraft of the present invention. In the illustrative embodiment, the inventive system is adapted for use with a satellite having a payload upon which attitude control elements are mounted. A single degree of freedom joint connects the bus to the payload such that the bus rotates about a single degree of freedom relative to the payload. An actuator and a control circuit therefore are included for rotating the payload relative to the bus.
In a specific embodiment, the bus is connected to the payload via a single degree of freedom rotational joint (often called a Bapta joint). The bus is driven to cancel any momentum of the payload about a first axis. This yields a spacecraft with zero net momentum. An additional controller is included to orient the spacecraft via the application of internally generated spacecraft forces or moments to control the spacecraft in the two axes orthogonal to the first axis.
In specific embodiment, the satellite includes a bus section and a payload section. The spacecraft bus serves as a storage section and accommodates the attitude control elements. It is foreseen that the bus will have a mass moment of inertia on the same order as the moment of inertia of the payload. The first control circuit controls the orientation of the bus relative to the payload via the Bapta joint. The first control circuit is a spin controller that computes an actuator control signal that drives the Bapta joint actuator and thereby spins the payload. The first control circuit further includes a rate detector for determining the payload angular rate relative to the bus about each of three axes. In the present specific embodiment, the rate detector includes a gyroscope sensor package and a tachometer in communication with the spin controller. The gyroscope sensor package provides a rate signal to the computer that is representative of the spacecraft momentum. The computer runs software to generate the spin control signal in response to the receipt of the rate signal from the gyroscope and tachometer sensor packages.
In addition to the spin axis control circuit, two additional control circuits are employed to control the two attitude angles perpendicular to the spin axis. The type of control circuits required is application-specific and may be determined by one skilled in the art to meet the needs of a given application.
In one embodiment, the additional controllers control a first reaction wheel having an axis of rotation perpendicular to the axis of rotation of a gimbal upon which the first wheel is mounted. The gimbal has an axis of rotation approximately parallel to the axis of rotation of the Bapta mechanism. The control circuits use angular position and rate measurements from sensors mounted on the gimbal and reaction wheel as well as angular rate signals from a gyroscope sensor package to generate torque signals via a computer to torque motors mounted on the gimbal and reaction wheel.
In a second embodiment, the additional controller includes a first reaction wheel and a second reaction wheel. The first and second reaction wheels are rigidly mounted to the spacecraft bus and are free to spin about first and second mutually perpendicular axes respectively. The first and second reaction wheels are selectively spun via first and second actuators in response to first and second steering control signals, respectively. The control signals are generated by a controller. The controller is in communication with a gyroscope sensor package, a star tracker, and tachometers mounted on the wheels, as is common in the art.
In a third embodiment, the additional controller includes a reaction mass that is mounted to the spacecraft via a flexure suspension. A first force actuator applies a first force to the reaction mass to facilitate spacecraft orientation. A second force actuator applies a second force to the reaction mass to facilitate spacecraft orientation. The first and second force actuators are voice coil actuators that selectively produce first and second forces, respectively, in response to the receipt of steering control signals. These actuators are placed so that the spacecraft can be steered in the two axes orthogonal to the spin axis. Proximity sensors measure the angle of the reaction mass relative to the spacecraft. Spacecraft steering torque commands are then generated by software running on a computer that is in communication with a gyroscope sensor package, star tracker, and the proximity sensors.
The unique design of the present invention is facilitated by the use of the counter-rotating bus and accompanying Bapta mechanism to cancel spacecraft momentum. Due to the mass characteristics of the bus, typically only relatively slow rotation of the bus is required to cancel spacecraft momentum. The slow rotation results in a minimum of spacecraft vibration. Any resulting vibrations are typically low frequency vibrations that are easily removed by controllers well known in the art. This obviates the need for an expensive momentum wheel isolation system and results in a more cost-effective spacecraft. Furthermore, the bus of the Bapta mechanism is useable as a storage compartment, making efficient use of use of spacecraft space and weight.